1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with total cooling of the entire airfoil.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
In the prior art, an airfoil leading edge is cooled with backside impingement cooling in combination with a showerhead arrangement of film cooling holes along with pressure and suction side film cooling (see FIG. 4). All leading edge region film cooling rows are supplied with cooling air from a common impingement cavity and discharge at various gas side pressures. As a result of this design, cooling flow distribution and pressure ratio across all of the leading edge region film cooling holes are both predetermined by the impingement cavity pressure. Also, the standard film cooling holes pass straight through the airfoil wall at a constant diameter and exit at an angle to the surface of the airfoil wall. Some of the coolant is subsequently ejected directly into the mainstream gas flow causing turbulence, coolant dilution, and a loss of downstream film cooling effectiveness. Further, the film cooling hole breakout on the airfoil leading edge surface many not achieve an optimum film coverage in a blade cooling application. The sidewall for the impingement cavity is cooled with a low heat transfer coefficient recirculation vortex created by the impingement jet. The cooling air supply cavity requires a reduction in the cross sectional flow area in the direction of the cooling air flow or through the flow Mach number in order to maintain adequate heat transfer capability as the cooling air is bled off from the cavity.